Multi-walled airfoil core

ABSTRACT

An airfoil core includes a first core portion that has a hybrid skin core, a tip flag core, and a trailing edge core. A second core portion has a serpentine core and a leading edge core.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a divisional of U.S. application Ser. No.16/202,483, filed Nov. 28, 2018 and now issued as U.S. Pat. No.11,015,457, which claims priority to U.S. Provisional Application No.62/741,322 filed Oct. 4, 2018 and claims priority to U.S. ProvisionalApplication No. 62/739,657, which was filed on Oct. 1, 2018.

BACKGROUND

This application relates to a multi-walled airfoil core and a method forproducing a multi-walled airfoil.

Gas turbine engines typically include a compressor section, a combustorsection, and a turbine section. In general, during operation, air ispressurized in the compressor section and is mixed with fuel and burnedin the combustor section to generate hot combustion gases. The hotcombustion gases flow through the turbine section, which extracts energyfrom the hot combustion gases to power the compressor section and othergas turbine engine loads.

Due to exposure to hot combustion gases, numerous components of a gasturbine engine, such as turbine blades, may include cooling schemes thatcirculate airflow to cool the component during engine operation. Thermalenergy is transferred from the component to the airflow as the airflowcirculates through the cooling scheme to cool the component.

Gas turbine engine airfoils, such as turbine blades and turbine vanes,can be fabricated by investment casting. For instance, in investmentcasting, a ceramic or refractory metal core is arranged in a mold andcoated with a wax material, which provides a desired shape. The waxmaterial is then coated with another material, such as a metallic orceramic slurry that is hardened into a shell. The wax is melted out ofthe shell and molten metal is then poured into the remaining cavity. Themetal solidifies to form the airfoil. The core is then removed, leavinginternal passages within the airfoil. Typically, the passages are usedfor cooling the airfoil.

SUMMARY

In one exemplary embodiment, a method of forming an airfoil, includesforming a hybrid skin core, a tip flag core, and a trailing edge core.The hybrid skin core, tip flag core, and trailing edge core areconnected to form a first core portion. A leading edge core and aserpentine core are formed. The first core portion, the leading edgecore, and the serpentine core are assembled together to form an airfoilcore. An airfoil is formed around the airfoil core.

In a further embodiment of any of the above, the airfoil is formed byinvestment casting.

In a further embodiment of any of the above, the hybrid skin core, tipflag core, trailing edge core, leading edge core, and serpentine coreare injection molded.

In a further embodiment of any of the above, the trailing edge core andthe hybrid skin core are formed concurrently in a first die.

In a further embodiment of any of the above, the tip flag core, trailingedge core, and hybrid skin core are formed concurrently in the firstdie.

In a further embodiment of any of the above, the leading edge core andthe serpentine core are formed concurrently in a second die.

In a further embodiment of any of the above, the leading edge core andthe serpentine core are connected to form a second core portion.

In a further embodiment of any of the above, a rod is inserted into theleading edge core before forming the airfoil.

In a further embodiment of any of the above, the rod extends to afeature radially outward of tip flag core.

In a further embodiment of any of the above, the rod is removed afterforming the airfoil to form a port near a leading edge of the airfoil.

In a further embodiment of any of the above, assembling the airfoil corecomprises inserting a rod into the serpentine core and the tip flagcore.

In a further embodiment of any of the above, the airfoil core containsone of ceramic, alumina, silica, and a metallic alloy.

In another exemplary embodiment, an airfoil core includes a first coreportion that has a hybrid skin core, a tip flag core, and a trailingedge core. A second core portion has a serpentine core and a leadingedge core.

In a further embodiment of any of the above, the hybrid skin coreportion, tip flag core and trailing edge core are injection molded as asingle structure.

In a further embodiment of any of the above, the serpentine core and theleading edge core are injection molded as a second single structure.

In a further embodiment of any of the above, the hybrid skin core isarranged along a suction side of the airfoil core. The serpentine coreis arranged along a pressure side of the airfoil core. The tip flag coreextends axially from the hybrid skin core to a trailing end of theairfoil core radially outward of at least a portion of the serpentinecore.

In a further embodiment of any of the above, a rod connects theserpentine core and the tip flag core.

In a further embodiment of any of the above, the hybrid skin core andtrailing edge core are connected at an inlet.

In a further embodiment of any of the above, a rod connects the trailingedge core and the tip flag core.

In a further embodiment of any of the above, the trailing edge core andthe tip flag core are connected via core material beyond the airfoiltrailing edge.

In another exemplary embodiment, an airfoil includes a body that extendsradially from a platform to a tip, axially from a leading edge to atrailing edge, and has a thickness between a pressure side and a suctionside. A hybrid skin cooling passage is arranged along the suction side.A tip flag cooling passage extends along the tip between the hybrid skincooling passage and the trailing edge and is in fluid communication withthe hybrid skin cooling passage. A trailing edge cooling passage extendsfrom the platform to the tip flag cooling passage.

In a further embodiment of any of the above, the serpentine coolingpassage has a turbulator on one passage surface.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2A is a perspective view of an airfoil.

FIG. 2B is a plan view of the airfoil illustrating directionalreferences.

FIG. 3A shows an airfoil core from the pressure side.

FIG. 3B shows an airfoil core from the suction side.

FIG. 4 shows a view of an airfoil core.

FIG. 5A shows a portion of an airfoil core from the pressure side.

FIG. 5B shows a portion of an airfoil core from the suction side.

FIG. 6A shows a portion of an airfoil core from the pressure side.

FIG. 6B shows a portion of an airfoil core from the suction side.

FIG. 7A shows a portion of an airfoil core from the pressure side.

FIG. 7B shows a portion of an airfoil core.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a nacelle15, and also drives air along a core flow path C for compression andcommunication into the combustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith two-spool turbofans as the teachings may be applied to other typesof turbine engines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis X relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to the fan 42through a speed change mechanism, which in exemplary gas turbine engine20 is illustrated as a geared architecture 48 to drive a fan 42 at alower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high) pressure turbine 54. Acombustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 may be arranged generallybetween the high pressure turbine 54 and the low pressure turbine 46.The mid-turbine frame 57 further supports bearing systems 38 in theturbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis X which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor, or aftof the combustor section 26 or even aft of turbine section 28, and fan42 may be positioned forward or aft of the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1 and less than about 5:1. Itshould be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine and that thepresent invention is applicable to other gas turbine engines includingdirect drive turbofans, having single or multiple fan streams.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

The turbine section 28 includes a rotor having one or more blades orairfoils that are rotatable about the engine axis X in the core flowpath C. FIG. 2A illustrates an example turbine blade 64. A root 74 ofeach turbine blade 64 is mounted to a rotor disk. The turbine blade 64includes a platform 76, which provides the inner flow path, supported bythe root 74. An airfoil 78 extends in a radial direction R (shown inFIG. 2B) from the platform 76 to a tip 80. The airfoil 78 providesleading and trailing edges 82, 84. The tip 80 is arranged adjacent to ablade outer air seal (not shown).

FIGS. 2A and 2B somewhat schematically illustrate an exterior airfoilsurface 79 extending in a chordwise direction C from a leading edge 82to a trailing edge 84. The airfoil 78 is provided between pressure(typically concave) and suction (typically convex) walls 86, 88 in anairfoil thickness direction T, which is generally perpendicular to thechordwise direction C. Multiple turbine blades 64 are arrangedcircumferentially in a circumferential direction A. The airfoil 78extends from the platform 76 in the radial direction R, or spanwise, tothe tip 80. It should be understood that the airfoil 78 may be utilizedin other parts of the gas turbine engine 20, or in other gas turbineengines.

The airfoil 78 is formed by investment casting a metallic alloy, such asa nickel alloy, or other material in a mold around an airfoil core 100,shown in FIGS. 3A-3B. As can be appreciated, the core 100 includessections that correspond to portions of the cooling passages of theairfoil 78. FIG. 3A shows the airfoil core 100 from the pressure side86, while FIG. 3B shows the airfoil core 100 from the suction side 88.The example airfoil core 100 includes a hybrid skin core 102, a tip flagcore 104, a trailing edge core 106, a serpentine core 108, and a leadingedge core 110.

The airfoil core 100 may be fabricated using variety of manufacturingmethods and used in conjunction with conventional investment castingprocesses. Core manufacturing methods may include, but are not limitedto conventional core die tooling, injection molding, flexible tooling,fugitive core, lithographic tooling, and/or advanced additivemanufacturing processes to directly fabricate integral ceramic cores forsingle and multi-walled cooling design concepts. Alternatively, theairfoil cooling configuration may be directly fabricated using laserpowder bed metal fusion additive manufacturing processes such as directmetal laser sintering (DMLS) or Selective Laser Sintering (SLS) methods.The core 100 is injection molded from a material that contains ceramic,silica or alumina aggregate and/or elements of a metal alloy material,for example. Solid structures define the core 100, which then producevoid structures in the airfoil 78, after the core material is leachedout of the metal airfoil as part of the investment casting process.These void structures make up a cooling passage network. Conversely,void structures of the core 100 produce solid structures in the airfoil78. The core 100 may have several internal geometric features definingthe internal core passages, cooling and heat transfer augmentationfeatures, (i.e.; trip strips, crossovers, impingement rib features), andcore producibility features, (i.e.; core printouts, chaplets, and corebumpers).

The hybrid skin core 102 forms a hybrid skin cooling passage thatextends radially and is provided in a thickness direction between thecore cooling passages and the suction side 88 of the airfoil 78. Thehybrid skin cooling passage provides double wall cooling, which improveslocal thermal cooling effectiveness of the turbine blade 64. In someembodiments, film cooling holes may be drilled into the hybrid skincooling passage to create a thin film boundary layer that protects theairfoil 78 from hot gases in the core flow path C. The hybrid skincooling passage receives cooling fluid from a cooling source at thehybrid skin core inlet 130. The hybrid skin cooling passages, locatedimmediately adjacent to the airfoil suction side leading edge, arepositioned to shield the leading edge feed cavity 110 as well as shieldother cooling passages, such as a serpentine cooling passage on thepressure side formed by the serpentine core 108. This shielding reducescooling air heat pickup of the airflow in the leading edge feed passage110 and the serpentine cooling passages by reducing the net heat fluxemanating from the hot external airfoil wall along the suction sidesurfaces immediately adjacent to the leading edge feed passage 110 andthe serpentine cooling passages. The reduction in cooling airtemperature heat pickup enables a more effective thermal cooling designconfiguration by maximizing the potential temperature gradient betweenthe external gas temperature and the colder internal cooling airtemperature. The reduction in cooling air heat pickup also lowers thetemperature of the local insulating boundary layer of the film coolingair, which thereby suppresses the local external heat flux andsubsequent external hot wall metal temperatures.

In some embodiments, the shielding provided by the suction side hybridskin core cavities may allow for the elimination of some turbulatorfeatures. For example, suction side internal trip strip features alongthe leading edge cooling passage 110 and serpentine cooling passagesimmediately adjacent to the cold internal wall formed between thesuction side hybrid skin cooling passages, and the leading edge coolingpassage 110 and the serpentine cooling passages may not be necessary.The additional convective heat transfer augmentation from trip strips orturbulators is not required along the internal cooling passage surfacesimmediately adjacent to the cold internal wall. Therefore, the overallpressure drop within the leading edge feed passage 110 and theserpentine cooling circuit may be significantly reduced due to the lowerfriction losses associated with cooling passages having internal tripstrips features present along only one of cooling passage surfacesimmediately adjacent to the external airfoil wall pressure side 86.

The tip flag core 104 forms a tip flag cooling passage extending fromthe hybrid skin core cooling passage to the trailing edge 84 near theairfoil tip 80. The tip flag core 104 may be joined to or integratedwith the hybrid skin core 102. This allows the cooling air from thehybrid skin passage to be reused to cool the tip flag cooling passage.The predominately axially flowing tip flag cooling passage is arrangedto maximize the internal convective heat transfer and provideuninterrupted access for optimization of tip film cooling holes,distributed along the pressure side tip surface, a pressure side tipshelf, and/or a tip squealer pocket. The ability to tailor the tip filmcooling enables the local thermal cooling effectiveness of the airfoiltip 80 to be significantly improved by more efficiently reducing localtip operating metal temperatures and increasing tip durability oxidationand thermal mechanical fatigue (TMF) capability. The improvement inairfoil tip durability ensures that a minimum tip clearance between theairfoil tip 80 and the blade outer air seal is maintained throughout thelife cycle of engine operation. The ability to minimize tip clearancebetween the airfoil tip 80 and a blade outer air seal is necessary toensure the optimum performance characteristics of the turbine areretained.

The trailing edge core 106 forms a trailing edge cooling passageextending radially between the pressure and suction sides 86, 88, andextending axially to the trailing edge 84. Cooling air enters thetrailing edge cooling passage via the trailing edge inlet 132. In someembodiments, the hybrid skin core 102 is connected to the trailing edgecore 106 and the tip flag core 104. The hybrid skin core 102 andtrailing edge core 106 may be connected at the tip flag core 104 and/orat the inlets 130, 132 and/or at the exit of the tip flag exit 112 andthe exit of the trailing edge 84, external to the airfoil 78. In oneembodiment, additional stock may be added to extend core material of thetip flag exit 112 and the exit of the trailing edge 84 in order toconnect the two core passages at a location external to the airfoil 78trailing edge geometry (shown in FIG. 7A). This arrangement may helpminimize the relative displacement between the tip flag core 104 andtrailing edge core 106. Joining the suction side hybrid skin core 102,the tip flag core 104 and the trailing edge 84 during the core injectionand manufacturing process may significantly improve casting processcapability by improving internal and external wall control, relativecore displacement, and core true position tolerance during wax injectionand subsequent metal pour operations as part of the investment castingprocess.

The serpentine core 108 forms a serpentine cooling passage arrangedbetween the hybrid skin cooling passage and the pressure side 86 of theairfoil 78. The example serpentine core 108 provides “up” passages 120,122 fluidly joined by a “down” passage 124. Each of the passages 120,122, 124 extends in a generally radial direction. In an embodiment,cooling air flows forward. That is, the cooling air enters theserpentine core passage at an inlet 134, and flows upstream towards theleading edge 82. The serpentine cooling passage may be co-flowing orcounter-flowing, for example. At least one of the passages 120, 122, 124is shielded by the hybrid skin cavity on the suction side 88. In someembodiments, at least one passage 120, 122, 124 extends into a pressureside tip shelf and/or a recessed tip squealer pocket. As shown in FIGS.3A and 4, the passages 120, 122 are joined radially inward of the tip80. This permits the serpentine core 108 to fit radially inward of thetip flag core 104. In other words, the tip turn is reduced in radialheight such that the turn is recessed radially inward of the tip flagcore 104, minimizing the overall pressure loss and cooling air heatpickup due to the reduced streamwise length of the serpentine coolingpassage. The reduction in pressure loss may be further improved byincorporating trip strip or turbulator features 123 (shown schematicallyin FIG. 6A) in the passages 122, 124 only along the pressure sidesurface of the serpentine cooling passage. The opposite suction sidecooling passage surface does not require enhanced internal convectiveheat transfer since these surfaces of cooling passages 124 and 122 areimmediately adjacent to a cold internal wall that is formed between thehybrid skin core 102 and the serpentine core 108. Eliminating the tripstrips on the inner wall results in a smaller pressure loss, whichallows the passages to meet the minimum backflow pressure, even underworst case tolerances. The pressure loss within the serpentine coolingpassage is significantly diminished due to both the reduced streamwiseL/Dh (Length/Hydraulic Diameter) of the serpentine cooling passage, andthe one wall trip strip passages 124 and 122. The reduction in pressureloss within the serpentine passage enables a higher internal coolingpressure to be retained which may increase the pressure ratio acrossfilm cooling holes emanating from the serpentine cooling passage andincrease the nominal and minimum back flow margin (BFM) required forachieving effective film cooling characteristics. In some embodiments,the passages 120, 122, 124 may include pedestals, crossovers, and/orribs, depending upon the particular core design.

Additionally, the last “up” passage 122 of the serpentine coolingpassage is designed to “print-out” along the airfoil tip 80 via a port144, either within the pressure side tip shelf and/or within the tipsquealer pocket. The “print-out” port 144 of the last “up” passage 122of serpentine passage may be used to purge residual ceramic corematerial contained within the cast airfoil after alloy metal pour andsolidification. The port 144 is a purge feature that may be fabricateddirectly from core material during the core injection process.Alternatively the last “up” passage 122 may comprise of a “solid rod”feature 140 (shown in FIGS. 7A-7B) around which ceramic core material isinjected. In this instance the solid rod feature 140 may be made ofquartz, alumina, platinum, or any other high melt temperature material.The solid rod 140 enables improved wall control and positioning of thelast “up” passage 122, while also providing a method to stabilize theserpentine 108 cooling circuit passages and/or core section 142 forminga “plenum-like” geometry feature. In some embodiments, the port 144 isformed by the rod feature 140.

The leading edge core 110 forms a leading edge cooling passage whichpurges cooling air at the leading edge 82. In some embodiments, thecooling air is purged into a pressure side tip shelf. The leading edgecooling passage is also shielded by the hybrid skin core passage 102,which reduces heat pickup to the leading edge cooling passage. In someembodiments, a plurality of cavities 138 are arranged at the leadingedge 82, and deliver film cooling to the leading edge 82. Cooling airenters the leading edge cooling passage via the inlet 136. Thearrangement of the hybrid skin passage improves thermal cooling acrossthe leading edge 82, which reduces the film cooling and improvesdurability of the airfoil 78.

A solid rod feature 140 (shown in FIGS. 7A-7B) may also be integratedinto the leading edgecore 110 which forms the leading edge coolingpassage and serves a similar purpose as it pertains to serving as aconduit for leaching core material out of the cast airfoil 78.Additionally, the solid rod feature 140 emanating from the leading edgecore 110 may be combined and joined with another solid rod feature 140emanating for the last “up” passage core 122. Each of the solid rodfeatures 140 may be integrally “joined” by a common ceramic core“plenum-like” geometry feature 142 which extends radially outboard ofthe outer diameter of the airfoil tip 80, in which both solid rods 140emanating from the leading edge core 110 and the serpentine core 108 areencapsulated. In this sense the two cores 110 and 108 are physicallyconstrained in order to improve core stability and relative movement,thereby improving core manufacturing and casting process capability.

Complex multi-walled blades are desirable in high performance turbinesbecause of their effective use of cooling air, but such multi-walledblades require complex and expensive manufacturing techniques. One knowntechnique includes forming multiple cores that must be individuallyassembled into a full core. This may result in increased tolerancesbetween the cores, increased cost for multiple core dies, and increasedtime of assembly for multiple cores. Another known core manufacturingtechnique involves a fugitive core process which requires multiple waxinserts for the ceramic injection process, and typically slows the rateof production and is more expensive and complex. The disclosed core 100permits simplified manufacturing of a multi-walled blade.

The disclosed airfoil core 100 combines the suction side hybrid skincore 102 with the trailing edge core 106. These cores 102, 106 areconnected at the trailing edge 84 via the tip flag 104 running from theskin core 102 to the trailing edge 84. The cores 102, 106 may also beconnected at the root outside of the finished casting part. Thiscombination of the hybrid skin core 102 and trailing edge core 106allows the cores to be produced in a single die, reducing the die count.The combination also reduces the number of separate cores that need tobe assembled into the final core 100. The connectivity reducestolerances between the connected cores and provides better overall andrelative core positioning control.

A method of forming the disclosed core 100 includes forming the trailingedge core 106 and hybrid skin core 102 in a single die, as shown in FIG.5A-5B. The hybrid skin core 102 is connected to the trailing edge core106 via the tip flag core 104. The skin core 102 may also be connectedto the trailing edge core 106 at the root.

As shown in FIGS. 6A-6B, the leading edge core 110 and serpentine core108 are formed. The leading edge core 110 and serpentine core 108 may beformed in a single die, or may be formed separately and then assembledtogether. The cores may be formed by injection molding. In anembodiment, each of the cores 102, 104, 106, 108, 110 are formed from arefractory metal core body material, silica, alumina, or another ceramiccore body material.

In some embodiments, the hybrid skin core 102, tip flag core 104, andtrailing edge core 106 are injection molded concurrently in a single dieto form a contiguous first core portion. In some embodiments, theleading edge core 110 and serpentine core 108 are injection moldedconcurrently in a single die to form a contiguous second core portion.In other embodiments, some of these components are injected separately.The first core portion and second core portion are assembled to form anairfoil core 100, or cores 102, 104, 106, 108, 110 are assembledtogether to form the airfoil core 100.

In some embodiments, additional structures are used to maintain thearrangement of the cores 102, 104, 106, 108, 110 in the airfoil core100. For example, a quartz or alumina rod 140 may extend from theserpentine core 108 or the leading edge core 110. In another example, arod 140 may also be used in the trailing edge core 106 and tip flag core104. A rod 140 may be used to join any two adjacent cores 102, 104, 106,108, 110. Such rods 140 provide additional stabilization between cores,preventing relative movement during the manufacturing process.

Once the core portions 102, 104, 106, 108, 110 are assembled into theairfoil core 100, the airfoil 78 is formed about the airfoil core 100.The airfoil core 100 is placed into a wax tool, which is used to createa wax pattern of the exterior shape of the airfoil 78. This wax tool mayform the airfoil surface 79, platform 76, and, root 74 (shown in FIG.2A). After several wax patterns are formed, these are placed into anarray, and ceramic slurry and particles are poured over the wax patternsto create a mold shell. Finally, the wax is melted out to leave a voidin the mold shell, which is later filled with a molten metallicmaterial, for example.

This method provides a quicker and simpler process of producing amulti-walled blade. The core arrangement may also provide better controlof the core portions. For example, the disclosed core 100 enablestighter control of the rib between the trailing edge core and tip flagcore, and better positional control of the hybrid skin core.

It should also be understood that although a particular componentarrangement is disclosed in the illustrated embodiments, otherarrangements will benefit herefrom. Although particular step sequencesare shown, described, and claimed, it should be understood that stepsmay be performed in any order, separated or combined unless otherwiseindicated and will still benefit from the embodiments of the presentinvention. Additionally it is important to note that any complexmulti-facetted resupply geometries that bridge centrally located mainbody cooling passages and peripherally located hybrid skin core coolingcavity passages can be created at any number of radial, circumferential,and/or tangential locations within an internal cooling configuration.The quantity, size, orientation, and location will be dictated by thenecessity to increase the local thermal cooling effectiveness andachieve the necessary thermal performance required to mitigate hotsection part cooling airflow requirements, as well as, meet part andmodule level durability life, stage efficiency, module, and overallengine cycle performance and mission weight fuel burn requirements.

Additionally, those of ordinary skill in this art should recognize thatcore fabrication methods utilizing fugitive cores, injection molded, andadditive manufacturing methods may be implemented to provide evengreater robustness and flexibility in the manufacturing of complexmulti-core/multi-wall design configurations similar to the final core100. These advanced core manufacturing processes eliminate therequirement and need for multiple complex multi-pull and/or single pullcore die tooling and core assembly. Additionally, geometric core passageand core passage cooling design geometry features (i.e.; trip strips,pin fins, pedestals, cross overs, impingement ribs, etc.) associatedwith conventional core die tooling are no longer limited. Core passageand cooling feature geometries constrained to specific shapes, sizes,and orientations, due to core manufacturing limitations associated withcore die tooling such as back-lock, die pull orientations, and corefinishing processes may be significantly reduced and/or eliminatedallowing more cooling design flexibility and optimization of integralmulti-wall core cooling configurations.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

The invention claimed is:
 1. An airfoil core, comprising: a first coreportion having a hybrid skin core, a tip flag core, and a trailing edgecore distributed in a chordwise direction between a leading edge of theairfoil core and a trailing edge of the airfoil core, wherein thetrailing edge core is adjacent to a trailing edge of the airfoil core,the hybrid skin core and the trailing edge core extend in a spanwisedirection, the tip flag core extends in the chordwise direction along atip of the airfoil core from the hybrid skin core to the trailing edgeof the airfoil core, the tip flag core extends in a thickness directionbetween a pressure side and a suction side of the airfoil core, and thehybrid skin core is arranged along the suction side of the airfoil core;and a second core portion having a serpentine core and a leading edgecore distributed in the chordwise direction, wherein the leading edgecore is adjacent to the leading edge of the airfoil core such that theleading edge core and the hybrid skin core are aligned in the chordwisedirection, and the serpentine core is arranged along the pressure sideof the airfoil core.
 2. The airfoil core of claim 1, wherein the hybridskin core, the tip flag core and the trailing edge core are injectionmolded as a single structure.
 3. The airfoil core of claim 1, whereinthe serpentine core and the leading edge core are injection molded as asecond single structure.
 4. The airfoil core of claim 1, wherein a rodconnects the serpentine core and the tip flag core.
 5. The airfoil coreof claim 1, wherein the hybrid skin core and the trailing edge core areconnected at an inlet.
 6. The airfoil core of claim 1, wherein a rodconnects the trailing edge core and the tip flag core.
 7. The airfoilcore of claim 1, wherein the trailing edge core and the tip flag coreare connected via core material along the trailing edge of the airfoilcore.
 8. The airfoil core of claim 1, wherein a rod extends from theleading edge core to a feature radially outward of the tip flag core. 9.The airfoil core of claim 8, wherein a second rod extends radiallyoutward of the serpentine core to the feature.
 10. The airfoil core ofclaim 1, wherein the airfoil core has a plurality of heat transferaugmentation features.
 11. The airfoil core of claim 1, wherein theserpentine core is arranged inward of the tip flag core relative to thespanwise direction.
 12. The airfoil core of claim 1, wherein the leadingedge core is configured to form a plurality of cavities spaced radiallyalong a leading edge.
 13. The airfoil core of claim 1, wherein: theserpentine core includes first, second and third serpentine sectionsdistributed in the chordwise direction to establish a serpentinegeometry, wherein each of the first, second and third serpentinesections extends in the spanwise direction, and wherein the secondserpentine section interconnects the first and third serpentine sectionsat respective bends at opposite ends of the second serpentine section;and the tip flag core includes first and second flag portions, the firstflag portion extends in the chordwise direction from the hybrid skincore at the tip of the airfoil core, the second flag portion extends inthe chordwise direction from the first flag portion to the trailing edgeof the airfoil core, the first flag portion and the second serpentinesection are aligned in the thickness direction between the pressure andsuction sides, the second flag portion and the third serpentine sectionare aligned in the chordwise direction, and the second flag portionextends in the thickness direction between the pressure and suctionsides such that the second flag portion is between the third serpentinesection and the tip of the airfoil core relative to the spanwisedirection.
 14. The airfoil core of claim 13, wherein the serpentine coreis arranged in the chordwise direction between the leading edge core andthe trailing edge core.
 15. The airfoil core of claim 14, wherein thefirst flag portion flares outwardly in the thickness direction from thehybrid skin core to the second flag portion to establish the pressureside of the airfoil core.
 16. The airfoil core of claim 15, wherein thesecond serpentine section and the first flag portion are aligned in thechordwise direction.
 17. The airfoil core of claim 16, wherein theleading edge core and the first serpentine section are spaced apart fromeach other in the chordwise direction.
 18. The airfoil core of claim 17,wherein: the hybrid skin core includes a first column of slots and asecond column of slots extending in the spanwise direction, and whereinthe first column of slots are spaced apart from the second column ofslots with respect to the chordwise direction; and the first column ofslots includes a first slot, the second column of slots includes asecond slot, and the first and second slots are aligned in the spanwisedirection with the tip flag core.
 19. The airfoil core of claim 18,wherein: an inlet portion of the hybrid skin core and an inlet portionof the trailing edge core are joined together at a position inward ofthe first and second columns of slots relative to the spanwise directionto establish the first core portion; and an inlet portion of theserpentine core and an inlet portion of the lead edge core are joinedtogether at a position inward of the respective bends of the serpentinecore relative to the spanwise direction to establish the second coreportion.
 20. The airfoil core of claim 16, wherein the first serpentinesection extends in the spanwise direction to the tip of the airfoil coresuch that the first serpentine section is arranged in the chordwisedirection between the leading edge core and the first flag portion.